The present invention relates to a wing arrangement for an aircraft comprising a wing having a base section and a tip section, the base section having a first end portion and a second end portion, the tip section having a third end portion and a fourth end portion, wherein the first end portion is adapted to be coupled to the fuselage of the aircraft and wherein the second end portion and the third end portion are coupled with each other so that the tip section is pivotable with respect to the base section about a pivot axis.
The aspect ratio, i.e. the ratio of span to chord, of an aircraft wing is one factor influencing the efficiency of the aircraft during flight. Generally, an increase of the aspect ratio is associated with an increased efficiency during steady flight. Therefore, an increase of the wingspan of an aircraft is one factor to take into consideration when seeking to reduce fuel consumption. However, when elongating the wing of an existing family of aircraft, it may become necessary to adapt the aircraft family specific infrastructure, and airport fees may increase.
One possibility to increase the wingspan without having to adapt the aircraft family specific infrastructure and having to deal with increased airport fees, or to reduce airport fees for existing aircraft is to provide for a foldable wing construction which allows to pivotably move an outboard end portion of the wing between a deployed position, in which the wing has its normal flight configuration, and a stowed position, in which the wing has a folded configuration and the wingspan is decreased as compared to the deployed position.
The present invention may be embodied as a wing arrangement having a foldable wing which is safe and reliable in operation and has at the same time a simple construction.
A wing arrangement according to an embodiment of the present invention comprises a wing having a base section and a tip section, the base section having a first end portion and a second end portion, the tip section having a third end portion and a fourth end portion, wherein the first end portion is adapted to be coupled to the fuselage of the aircraft, wherein the second end portion and the third end portion are coupled with each other so that the tip section is pivotable with respect to the base section about a pivot axis. In particular, a pivotal movement of the tip section relative to the base section between a stowed position and a deployed position is enabled, wherein the spanwise length of the wing is larger in the deployed position than in the stowed position. In other words, the length of the entire wing measured parallel to the y-axis of the aircraft to which the wing arrangement is coupled, is larger in the deployed position than in the stowed position.
Thus, the wing arrangement is formed of two sections, namely the base section intended to be secured to the fuselage of the aircraft and the tip section pivotably connected to the distal end of the base section. Further, in one embodiment the pivot axis may extend in chord direction of the wing so that the tip section pivots about an axis which is arranged horizontally or essentially parallel to the x-axis of the aircraft, and in the stowed position the tip section extends vertically. In another embodiment the pivot axis may extend nearly parallel to the z-axis of the aircraft on which the wing arrangement of the present invention is mounted. In particular, it can be tilted by 15° with respect to the vertical direction or z-axis so that when the tip section pivots rearwards or forwards from the deployed position in which the tip section is aligned with the longitudinal axis of the base section, to the stowed position, the distal end of the tip section moves upwards.
Furthermore, the wing arrangement of the present invention comprises an actuating arrangement having an actuator which is coupled to the base section and the tip section and which is operable to effect the pivotal movement of the tip section relative to the base section between the stowed position and the deployed position. The actuator has a rotatable actuator input shaft which is rotatably supported on the base section, and the actuator is configured such that rotationally driving the actuator input shaft effects a pivotal movement of the tip section relative to the base section about the pivot axis between the deployed and stowed positions.
In addition, the actuating arrangement comprises a differential gearbox having a rotatable output shaft, a first rotatable input shaft and a second rotatable input shaft, and the first input shaft, the second input shaft and the output shaft are coupled such that rotary power applied to the first and/or second input shaft is transmitted to the output shaft, wherein the first and second input shafts are capable of being simultaneously rotatingly driven with different rotational speeds. Thus, the differential gearbox is configured as a common differential gear assembly which is capable of transmitting rotary power input via the first and second input shafts to the output shaft wherein the rotational power is combined.
Finally, the actuating arrangement comprises a first motor and a second motor each having a rotationally driven output shaft, wherein the first input shaft of the differential gearbox is coupled to the output shaft of the first motor and the second input shaft of the differential gearbox is coupled to the output shaft of the second motor. Hence, the input shafts of the differential gearbox can rotationally be driven by the motors. The output shaft of the differential gearbox is coupled to the actuator input shaft so that it is rotationally driven by the output shaft of the differential gearbox.
The assembly of the first and second motor which are both coupled with the inputs of the differential gearbox so that the output thereof drives the actuator allows to employ two independently driven motor to supply torque to the actuator. Hence, when one of the motors or the respective power supply fails, the actuator can still be operated by the other motor.
In one embodiment, the pivotal movement of the tip section is effected by operating only one of the motors at the same time whereas the other motor is not supplied with power so that it remains passive. In this embodiment the actuating arrangement is operated in the passive/active mode.
Further, the use of the above-described differential gearbox also allows for a so-called active/active mode of operation in which both motors are simultaneously operated to pivot the tip section so that the rotational power supplied by the motors is combined at the actuator input shaft to drive the latter. In a configuration employing the active/active mode smaller, less powerful motors can be used which leads to weight savings. However, the motors still have to be dimensioned such that operation of only one motor is sufficient to pivot the tip section even if this occurs with a lower speed compared to the situation in which both motors are operated.
In an embodiment the first input shaft of the differential gearbox is coupled to a first brake assembly which is operable to selectively allow or prevent rotation of the first input shaft. Further, the second input shaft of the differential gearbox is coupled to a second brake assembly which is operable to selectively allow or prevent rotation of the second input shaft. Such an arrangement is particularly useful when the drag torque of the motors coupled to the input shafts of the differential gear box is small so that if only one of the motors is operated the other would be reversely driven by the input shaft connected to it. In such case the respective brake assembly would be activated to prevent such reverse driving of the motors.
In another embodiment the first motor is a hydraulic motor and has a first hydraulic connection which is connected to a first connector being adapted to be coupled to a first hydraulic supply of the aircraft, the first hydraulic supply being capable of supplying a plurality of consumers in the aircraft. This allows to supply at least one of the motors for pivoting the tip section by a hydraulic system of the aircraft.
If the aircraft to which the wing arrangement of the present invention is coupled comprises a second supply system which is capable of supplying pressurized hydraulic fluid independent from the first hydraulic supply system, it is further preferred when the second motor is also a hydraulic motor and has a second hydraulic connection which is provided with a second connector being adapted to be coupled to the second hydraulic supply of the aircraft. In such case the redundancy of the actuation mechanism for pivoting the tip section relies upon the fact that the independently acting hydraulic systems used in the aircraft are employed as power sources.
As an alternative the second motor can be an electric motor, so that the first motor is hydraulically driven whereas the second motor is electrically driven. This also results in the required redundancy but avoids that further hydraulic conduits for connecting with the second hydraulic system of the aircraft have to be guided to the distal end of the base section.
Finally, both the first and the second motors may be electric motors which entirely avoids the necessity of hydraulic conduits being guided to the distal end of the base section of the wing.
The actuating arrangement includes a latching arrangement with a latching device which comprises a support fixed to one of the base section and the tip section, a latching element moveable relative to the support between a latching position and a release position, a first actuation element and a second actuation element. Further, the latching arrangement comprises an engagement element mounted on the other of the base section and the tip section, the latching device and the engagement element being configured such that when the tip section is in the deployed position and the latching element is in the latching position, the latching element engages with the engagement element so as to prevent the relative pivotal movement between the base section and the tip section, and when the latching element is in the release position, the latching element is disengaged from the engagement element, so that the pivotal movement between the base section and the tip section is enabled, wherein the actuation elements are coupled to the latching element and configured such that when at least one of the actuation elements is operated, the latching element is forced into the release position.
Thus, the latching device is configured to latch the tip section in the deployed position so as to prevent that the tip pivots away from the deployed position towards the stowed position. The tip section is released when the latching element is moved to the release position by means of the actuation elements. Further, by means of two actuation elements coupled to a sole latching element, it is ensured that if one of the actuation elements or its power supply fails, the latching element can still be operated by the other actuation element, so as to have a redundant system.
Similarly, the latching device and the engagement means may be further configured such that when the tip section is in the stowed position and the latching element is in the latching position, the latching element is in engagement with the engagement means so as to prevent a pivotal movement of the tip section away from the stowed position. Hence, the latching arrangement may also be able to latch the tip section in stowed position.
In a further embodiment, the first and the second actuation elements are hydraulically powered, and the first actuation element is connected to a first connector which is adapted to be coupled to a first hydraulic supply of the aircraft, the first hydraulic supply being capable of supplying a plurality of consumers in the aircraft. In addition, the second actuation element is connected to a second connector which is adapted to be coupled to a second hydraulic supply of the aircraft, the second hydraulic supply being also capable of supplying a plurality of consumers in the aircraft.
Thus, similar to the motors coupled with the differential gearbox, the actuation elements of the latching device may also be supplied by independently working hydraulic supplies or systems of the aircraft on which the wing arrangement is mounted. In particular, the connectors of the first motor and the first actuation element may be connected to the same first hydraulic system whereas the connectors of the second motor and the second actuation element may be coupled with the same second hydraulic system of the aircraft.
Alternatively, the first and the second actuation elements are hydraulically powered, the first actuation element is connected to a first connector which is adapted to be coupled to a hydraulic supply of the aircraft, the hydraulic supply being capable of supplying a plurality of consumers in the aircraft. Further, the actuating arrangement comprises a hydraulic pump which includes a rotatable input shaft and a hydraulic output, and an electric drive motor the output of which is coupled with the input shaft of the hydraulic pump. Finally, the hydraulic output of the hydraulic pump is hydraulically connected with the second actuation element.
Hence, in this embodiment one actuation element is supplied by a hydraulic supply of the aircraft whereas the other actuation element is supplied with pressurized hydraulic fluid provided by a hydraulic pump which is independently driven by an electric motor, so a sole hydraulic supply has to connected with elements at the distal end of the base section of the wing. Here, the redundancy is ensured by the electric motor combined with the hydraulic pump which are both operated when the hydraulic supply of the aircraft fails.
Moreover, both the first and the second actuation elements may be electrically powered, so that the latching device is independent from any hydraulic supply of the aircraft.
In one embodiment, the actuator comprises a threaded spindle rotatably supported and axially fixed on the base section, and a nut member which threadingly engages with the threaded spindle and which is mounted on the tip section in such a manner that it cannot rotate with respect to the longitudinal axis of the threaded spindle. Further, the actuator input shaft is coupled to the threaded spindle so that rotation of the actuator input shaft effects rotation of the spindle or the actuator input shaft is part of the threaded spindle. Thus, in this embodiment the actuator is formed as a jack screw arrangement which can be combined in a simple manner with a tip section which is pivotable with respect to a pivot axis that extends in chord direction of the wing, i.e. it is essentially parallel to the x-axis of the aircraft which axis essentially corresponds to the longitudinal axis of the fuselage.
In a further embodiment, a spindle locking device is interconnected between the actuator input shaft and the threaded spindle and which is configured such that the threaded spindle is brought into engagement with the second end portion so as to prevent rotation of the threaded spindle, when the threaded spindle transmits torque to the actuator input shaft, and that when torque is transmitted from the actuator input shaft to the threaded spindle, the threaded spindle is rotatable with respect to the second end portion. Such a so-called “No-Back Mechanism” is particularly advantageous when the afore-mentioned assembly with a threaded spindle is employed to pivot the tip section upwards about an essentially horizontal pivot axis. In this case the no-back mechanism prevents the tip section from pivoting from the stowed position back to the deployed position due to its own weight. Hence, it is not required that additional latching means are provided to latch the tip section in the stowed position.
In another embodiment, the actuator comprises a base member and an output member, wherein the output member is pivotably mounted on the base member so that the output member may pivot with respect to the base member about the pivot axis. The base member is fixed to the second end portion and the output member is fixed to the third end portion, i.e. the base member is secured to the base section of the wing, whereas the output member is secured to the pivotable tip section. A gear assembly is connected to the actuator input shaft and the output member, the gear assembly being configured such that rotationally driving the actuator input shaft with a first rotational speed results in a pivot movement of the output member relative to the input member about the pivot axis with a second rotational speed smaller than the first rotational speed. Such a geared rotary actuator can designed with small dimensions in the direction of the pivot axis, e.g. by employing planetary gear stages in the gear assembly, so that it can employed when the pivot axis about which the tip section pivots, extends in an essentially vertical direction, i.e. it is nearly parallel to z-axis of the aircraft and, as discussed above, preferably tilted by about 15° to the z-axis.
Finally, the actuating arrangement of the wing arrangement of the present invention may comprise a first end stop fixedly connected to the base section and a first abutment member fixed on the tip section, the first end stop and the first abutment member being configured such that in the deployed position the first end stop and the first abutment member abut on each other and prevent a further pivotal movement of the tip section relative to the base section in a direction opposite to the direction towards the stowed position.
In addition, a second end stop is fixed to the base section, and a second abutment member is fixed to the actuator input shaft, the second end stop and the second abutment member being positioned such that an abutment of the second end stop and the second abutment member prevents a further rotation of the actuator input shaft in a rotational direction that effects a pivotal movement of the tip section about the pivot axis towards the deployed position.
The positions of the first and second end stops and the first and second abutment members are chosen such that when the first end stop abuts on the first abutment member, the second end stop is spaced from the second abutment member so that a predetermined amount of rotational movement of the actuator input shaft in that rotational direction which effects a pivotal movement of the tip section towards the deployed position is required to bring the second end stop into abutment with the second abutment member.
Thus, due to the fact that the first end stop abuts on the first abutment member before there is an abutment between the second end stop and the second abutment member, it is possible to apply a predetermined bias or force which presses the tip section against the first end stop and prevents that due to play and elasticity in the drive train formed by the motors and the actuator the tip section may pivot backwards away from the first end stop. This bias or force is obtained by also bringing the second end stop member into contact with the second abutment member.
So, the tip section is kept in position on the first abutment member which is particularly desirable, when a latching device is employed the latching element of which can only brought into engagement with the respective engagement member, when the tip section is exactly in the deployed position.
However, as there is a well defined distance between the abutment positions the bias or force applied to the tip section and the first end stop and the first abutment member, respectively, is limited. This is particular useful, because the torque at the output of the actuator, in particular when comprising a reduction gear assembly, is much higher compared to the torque applied to actuator input shaft, so that a slightly excessive torque at the input shaft may result in much higher forces applied to the first end stop and the first abutment member which would structurally have to be accounted for when there was no additional end stop assembly on the actuator input shaft. Such excessive torque can readily occur as it is often not possible to stop the motors coupled to the differential gearbox at a correct point in time and there is the additional uncertainty due to the inertia of the differential gear box.
Finally, the above embodiments may be included in an aircraft comprising a fuselage and at least one of the above-described wing arrangements.